A gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section and an exhaust section. In operation, air enters an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section through a hot gas path defined within the turbine section and then exhausted from the turbine section via the exhaust section.
In particular configurations, the turbine section includes, in serial flow order, a high pressure (HP) turbine and a low pressure (LP) turbine. The HP turbine and the LP turbine each include various rotatable turbine components, such as turbine rotor blades, rotor disks and retainers, and various stationary turbine components, such as stator vanes or nozzles, turbine shrouds and engine frames. The rotatable and stationary turbine components at least partially define the hot gas path through the turbine section. As the combustion gases flow through the hot gas path, thermal energy is transferred from the combustion gases to the rotatable turbine components and the stationary turbine components.
Nozzles utilized in gas turbine engines, and in particular HP turbine nozzles, are often arranged as an annular array of nozzle segments, with each nozzle segment including one or more nozzle vanes or airfoils extending between annular inner and outer bands which define the primary flowpath through the nozzles. Due to the operating temperatures within the gas turbine engine, it is generally desirable to utilize materials having a low coefficient of thermal expansion and high compression strength. Recently, for example, ceramic matrix composite (“CMC”) materials have been utilized to operate effectively in such adverse temperature and pressure conditions. These low-coefficient-of-thermal-expansion materials have higher temperature capability than similar metallic parts, so that, when operating at the higher operating temperatures, the engine is able to operate at a higher engine efficiency.
Despite their high temperature capabilities, it is often desirable to provide a flow of cooling medium to hot gas path components formed from CMC materials. For instance, CMC-based nozzle designs are known in which the nozzle airfoil defines two radially extending internal cavities, namely a larger forward cavity extending from the airfoil's leading edge towards its trailing edge and a smaller aft cavity positioned between the forward cavity and the trailing edge. The forward and aft internal cavities are separated from one another in a chordwise direction of the airfoil by a vertical rib that extends radially between the two cavities along the radial height of the airfoil. As such, each cavity is separately supplied with cooling medium from a source disposed radially outwardly from the radial outer end of the vertical rib.
Unfortunately, such conventional cooling arrangements exhibit certain disadvantages. For example, the vertical rib dividing the forward and aft cavities generally creates an area of high thermal stress within the airfoil. In addition, the design of the aft cavity within the airfoil is typically limited by the radius of the aft end of such cavity (e.g., the end located closest to the trailing edge of the airfoil). Specifically, as the aft end of the aft cavity is moved closer to the trailing edge of the airfoil, the radius at the aft end must be decreased, thereby increasing the likelihood of failure due to the internal stresses caused by the airfoil “ballooning” during operation of the gas turbine engine. As a result, the aft end of the aft cavity is spaced apart from the trailing edge of the airfoil by a greater chordwise distance than is typically desired for optimal cooling performance so as to prevent the occurrence of airfoil failures.
Accordingly, an improved cooling flow arrangement for an airfoil of a gas turbine nozzle that allows for a cooling medium to be supplied within the airfoil in closer proximity to the trailing edge without increasing the likelihood of a failure occurring would be welcomed within the technology.